1. Field of the Invention
The present invention relates to the fabrication of composite structural units, and more particularly to a method for fabricating stiffened composite panel structures possessing hollow, U-shaped stiffener elements, spars, and transversely extending shear ties, where the fabrication and curing of the composite panel structures is accomplished without the use of mandrels.
2. Discussion of the Known Prior Art
With the increasing use of composite materials in the manufacture of aircraft components, a number of problems have arisen in connection with the application of such materials to design principles developed for conventional aluminum or titanium construction. Most of these problems center about the difficulty in interconnecting one component to one another, and involve the use of mechanical fasteners.
The currently-known advanced composite materials do not readily lend themselves physically or economically to extensive use of mechanical fasteners nor to very complex designs. The inclusion of multiple details in a composite material component increases the cost of manufacture to unacceptable levels and makes the application of automated procedures difficult, if not impossible. Mechanical fasteners are not physically well-suited to composite material components, and their use with such components tends to detract from the advantages of the unidirectional properties of advanced composite fibers. These problems have created a need for developing new techniques for economically producing composite material components, especially such components that are required to be load resisting.
In recent years, there have been a number of proposals relating to the structure and manufacture of composite material components. Some of these proposals involve the use of thermosetting adhesives instead of mechanical fasteners.
Manufacturing processes that use thermosetting adhesives generally require the separate forming and curing of the components and a further heating process to set the adhesives. Such multiple step processes have the disadvantages of being time consuming and expensive to carry out.
Other processes involve the separate forming of elements of a componenet and the curing together of the elements to form the desired structure. A serious problem encountered in connection with known composite material components embodying both approaches, i.e., bonding by adhesives and bonding by curing, has been the tendency for one part of the component to be peeled away from an adjacent part of the componenet when the component is subjected to stress forces. Recent proposals for overcoming such peel tendencies have the disadvantage of being quite complicated and expensive to carry out.
Aircraft components which have been formed from composite materials are disclosed by U.S. Pat. No. 3,995,080 Cogburn et al. U.S. Pat. No. 3,995,081 to Fant et al. These patents disclose a composite material plicated structural beam with a fairly complex design that is described as being peel-resistant. Each of the elements of the beam is formed separately, and then the elements are assembled and cured together to form the beam. The use of a destructible mandrel in the curing process is described.
Integrally stiffened structural shapes, fabricated by the lamination and resin polymerization or curing of resinous, filament reinforced composite materials, have typically taken the form of J-section and I-section structures. These designs have proven to be the most readily producible, but with great difficulty. For one thing, too many tooling variables must be considered; for another, it is clearly not cost effective to produce a J-section integrally stiffened structure in one cure cycle.
Therefore, a great need has arisen for a method of readily producing stiffened, fiber-reinforced composite structures useful in the construction of integrally stiffened components for aerospace vehicles which are cost and labor efficient and which save time in the fabrication process.